The present invention relates to gas turbine engine compressors, and more particularly to capturing air from a compressor tip for auxiliary uses. More specifically, it pertains to using the captured air supply for the passive cooling of an auxiliary power unit.
Gas turbine engine powerplants are used in the vast majority of aircraft flying today. Most large commercial aircraft include an auxiliary power unit (APU), generally a small gas turbine engine, often mounted in the aft tail section of the aircraft, which provides electrical power and pressurized air for aircraft environmental control systems when the aircraft is on the ground, and is also used to start the main engines of the aircraft. APUs require external cooling and are lubricated by oil that is generally cooled by an air cooled oil heat exchanger.
Active cooling systems are most often employed to provide this cooling air, and are typically comprised of a fan used to push air through the oil cooler and across auxiliary power unit components. These fans are driven at high speeds by the APU through relatively complex shaft and gear assemblies. The mechanical complexity and high operating speeds of these fans increase the possibility of failure of the cooling system, which would eventually lead to APU shutdown. Active fan cooling systems therefore significantly reduce the reliability of an auxiliary power unit, and add considerable cost and weight. While various passive cooling systems exist, they often require ducting air from the exterior of the aircraft, and fail to be able to provide compressed air for other uses.
Various systems used to separate compressor airflow are known. U.S. Pat. No. 5,357,742 issued Oct. 25, 1994 to Miller, for example, discloses metering cooling air exhausted through a turbojet laminar flow nacelle system, to cool the core engine compartment. Air bled from the entry to the core engine compressor drives a turbocompressor pump which draws cooling air through the laminar flow nacelle system and into a manifold surrounding the engine. This system has the disadvantage of requiring a separate pump to provide the compressed cooling air.
Separating airflow from the exit of a centrifugal compressor is also known. In U.S. Pat. No. 2,696,074 issued Jan. 2, 1953 to Dolza, an engine and torque converter cooling system having a two stage impeller and an annular diffuser is disclosed. Air is diverted from the main air stream flow, into either impeller stage. One or both of the impeller stages can be engaged. Two separate diffuser inlet nozzles accept air from each impeller stage and feed two diffuser chambers, one intended to cool the torque converter and the other the engine. The inlet airflow to the impeller is separated from its inlet and is selectively directed to one or both impeller stage inlets.
Passive cooling solutions particularly for auxiliary power units are numerous. U.S. Pat. No. 6,092,360 issued Jul. 25, 2000 to Hoag et al., discloses an APU passive cooling system in which cooling air is drawn into the engine compartment through an opening located in the rear of the aircraft. An eductor mounted before the exhaust duct of the engine, draws compartment air through the oil cooler, which in turn draws atmospheric air in through the aft opening.
Therefore, while methods of auxiliary power unit oil cooling and compartment pressurization exist which eliminate active cooling systems, there is a need for an APU built-in passive cooling system capable of providing compressed air for cooling and other uses. While some attempts have been made to use compressors as a source of cooling air, none employ the engine core compressor for a cooling system that does not require additional ducting of cooling air from the exterior of the aircraft.
It is an object of the present invention to supply cool air from the compressor of a gas turbine engine to be used for a means other than power generation.
It is another object of the present invention to fulfil the cooling and compartment pressurization requirements of an auxiliary power unit in an aircraft.
Therefore, in accordance with the present invention, there is provided a gas turbine engine compressor, comprising: a rotor adapted to rotate about a central axis, the rotor having a hub and rotor blades extending radially from the hub; an annular compressor casing being concentric with said central axis and defining an outer wall; said rotor blades having tips wherein at least part of said tips are in close proximity with said outer wall, and said blades having end portions near said tips; said outer wall extending upstream of said rotor, permitting substantially unobstructed fluid flow communication between an exterior air source and said rotor; an annular shroud within said compressor casing and concentric with said central axis, extending downstream from said rotor; a first annular duct defined within said annular shroud; said annular shroud and said outer wall defining a second annular duct; said first duct permitting core fluid flow communication between said rotor and a compressor outlet; and said second duct adapted to supply air from at least said end portions of said blades for auxiliary use.